This invention relates to gas turbines and, more particularly, to a concept for efficiently cooling ultrahigh temperature turbine rotor blades.
It is well understood that gas turbine engine shaft horsepower and specific fuel consumption, which is the rate of fuel consumption per unit of power output, can be improved by increasing turbine inlet temperatures. However, current turbines are limited in inlet temperature by the physical properties of their materials. To permit turbines to operate at gas stream temperatures which are higher than the materials can normally tolerate, considerable effort has been devoted to the development of sophisticated methods of turbine cooling. In early turbine designs, cooling of high temperature components was limited to transferring heat to lower temperature parts by conduction, and air cooling was limited to passing relatively cool air across the face of the turbine rotor disks.
In order to take advantage of the potential performance improvements associated with even higher turbine inlet temperatures, modern turbine cooling technology utilizes hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in the 2000.degree. to 2300.degree. F. (1094.degree. to 1260.degree. C.) range. Various techniques have been devised to air cool these hollow blades and vanes. These incorporate three basic forms of air cooling, either singly or in combination, depending on the level of gas temperatures encountered and the degree of sophistication permissible. These basic forms of air cooling are known as convection, impingement and film cooling. U.S. Pat. Nos. 3,700,348 and 3,715,170, assigned to the assignee of the present invention, are excellent examples of advanced turbine air-cooling technology incorporating these basic air-cooling forms.
However, the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle. For example, probably the most popular turbine coolant today is air which is bled off of the compressor portion of the gas turbine engine and is routed to the hollow interior of the turbine blades. Typically, the work which has been done on this air by the compressor is partially lost to the cycle. Additionally, as the cooling air circulates throughout the turbine blade it picks up heat from the metallic blades or vanes. When this heated cooling air leaves the turbine blades, perhaps as a coolant film, this heat energy is lost to the cycle since the hot gases are normally mixed with the exhaust gases and ejected from an engine nozzle. It would be desirable, therefore, to have a cooling system wherein a medium other than compressor bleed air is used and wherein the heat extracted by the cooling medium is put back into the cycle in a useful and practical manner.
A partial solution to the foregoing problems has been the suggestion of closed-loop cooling systems for turbine blades which may or may not also incorporate the concept of regeneration or recuperation to recover lost thermal energy. One such cooling arrangement which has been proposed, for example, is that of U.S. Pat. No. 2,782,000. In that patent, a closed-system steam thermosiphon is used to cool the turbine blades, the thermosiphon principle being that by which a coolant is caused to circulate throughout the hollow bores of a turbine blade under the pumping action of centrifugal force due to the difference in density between the heated coolant (steam) exiting the blade and the coolant (steam or water) entering it. Each blade is provided with its own thermosiphon which is associated with a cooler or heat exchanger which, in turn, is cooled by a second cooling medium such as water or air. However, the difficulties in fabricating the system, particularly the proposed fin-type heat exchangers, would preclude its practical application in aircraft gas turbines. Furthermore, the system requires the use of water (which is not normally available in an aircraft gas turbine engine environment) or compressor bleed air (with the same disadvantages discussed above) for the secondary coolant.
Another arrangement utilizing the closed-loop thermosiphon principle is that taught in U.S. Pat. No. 2,778,601 wherein hollow turbine blades are connected, via radially extending passages through the turbine disk, to a common, manifolded, fluid reservoir comprising a hollow turbine shaft and a coannular hollow feed tube. The disadvantages of this arrangement include the necessity of providing a pair of coannular hollow members for a single turbine stage. Since modern gas turbofan engines already incorporate as many as three coannular shafts without closed-loop thermosiphon cooling, the complexities of adding additional shafting for this purpose would be formidable indeed. Also, since all of the blades are manifolded at a common supply, a leak in one blade would result in a coolant loss in all blades. Still further, the disk passages tend to degrade the disk structural integrity, an important consideration in aircraft gas turbine engine design, and tend to increase its cost.
U.S. Pat. No. 2,849,210 teaches a turbine using the closed-loop thermosiphon principle wherein the hollow interiors are fluidly connected to an annular condensing chamber near the turbine disk bore by a plurality of tubes extending down one side of the disk between a manifolded vaporizing chamber associated with the blades and the manifolded condensing chamber. However, the loss of coolant from one blade would result in a loss of cooling capability for all blades. Also, recommended coolants are water, sodium, sulphur, potassium, mercury and fluorocarbons, none of which are normally available in gas turbine engines. Thus, additional weight would be added merely for the coolant, and it must be remembered that weight is another important consideration in aircraft gas turbine engine design. In an alternative embodiment, the patent teaches that the external coolant may be fuel or some of the compressed air which is fed to the engine combustor, in which instances at least a portion of the heat extracted from the turbine buckets would be returned to the gas turbine engine as usable energy. However, cooling a turbine blade by routing fuel through its hollow interior presents a potential fire hazard, and the use of compressor discharge air has the inherent disacvantages described above which the closed-loop system is intended to avoid.
Yet another cooling arrangement, for example, is that proposed in U.S. Pat. No. 2,883,151. In that patent the hollow turbine blades are again cooled by the closed-loop steam thermosiphon principle, with the rotor blade interiors communicating with longitudinally extending, coannular passages within the rotatable turbine shaft via radially extending passages through the turbine disc. The thermosiphon, in turn, is cooled by fuel circulating through another pair of coannular passages in a stationary stub shaft within the rotating shaft. Heat transfer is by convection between the rotating and stationary shafts. The difficulty of fabricating double coannular shafts is not insignificant. Other disadvantages which have been mentioned with regard to the previous systems include the manifolding of all blades to a common coolant supply, the proximity of fuel to the rotating hot turbine, and the degradation of integrity due to the radially extending bores therein.
One more arrangement, for example, is that of U.S. Pat. No. 3,756,020, wherein the turbine of a regenerative engine is cooled by a closed-system thermosiphon which passes through heat exchangers. These, in turn, are cooled by either fuel or air which are then routed to the combustor. However, the individual blade cooling circuits pass through the disc and are manifolded at the heat exchanger, thereby presenting disadvantages enumerated above.
It will therefore be appreciated that although the concepts of turbine cooling by the thermosiphon principle and regenerative engines are not new per se, a need exists for applying these concepts to an aircraft gas turbine engine in an efficient and reliable manner.